System and method to attach and remove space vehicles

ABSTRACT

A system and method for installing, deploying, and recovering a plurality of spacecraft that provides an ease of use and structural stability, and facilitates a standardization of spacecraft design. In embodiments of this invention, threaded rods are arranged orthogonal to a surface of a baseplate, and each spacecraft includes a coupling mechanism that selectively engages or disengages each threaded rod. Each spacecraft is added to the stack by engaging its coupling mechanism and rotating the threaded rods while the preceding spacecraft on the stack disengage their coupling mechanisms, thereby enabling the spacecraft to travel along the threaded rods toward the baseplate. When all of the spacecraft are added to the stack, the stack is preloaded by rotating the treaded rods into a terminator component at the top of the stack while the coupling mechanisms in all of the spacecraft are disengaged. Spacecraft are deployed by reversing the process.

This application is a Continuation of U.S. patent application Ser. No.16/511,864, which claims the benefit of U.S. Provisional PatentApplication 62/698,380, filed 16 Jul. 2018.

BACKGROUND AND SUMMARY OF THE INVENTION

This invention relates to the field of space craft, and in particular toa system and method for attaching multiple space vehicles to a launchvehicle, deploying the space vehicles, and recovering other spacevehicles.

To reduce the cost of space vehicle deployment, and in particular smallspacecraft (e.g. under 500 pounds), the U.S. Department of Defence(DoD), NASA, other government agencies, commercial companies, andseveral universities developed the EELV (Evolved Expendable LaunchVehicle) Secondary Payload Adapter (ESPA), which enables multiple smallspacecraft to be launched from ATLAS V, Falcon 9, and other launchvehicles. Other standards for multi-satellite payloads are alsoavailable, such as Loadpath's Cubestack, SSO-A's Multi Payload Carrierand Hub, and others, each with particular advantages and disadvantages.For ease of explanation and understanding, the EELV-ESPA standard isused herein to provide a contrast to embodiments of this invention.

FIGS. 1A-1C illustrate an example EELV Secondary Payload Adapter. Theadapter 100 is cylindrical, and includes six ports 110, each with anattachment flange 120. Each spacecraft 150 has a mating attachmentflange 130. Each spacecraft 150 is coupled to the adapter 100 bycoupling the attachment flanges 120, 130. This coupling may beaccomplished using a motorized LightBand 140 (U.S. Pat. Nos. 6,227,493;6,343,770; 6,390,416) from Planetary Systems, Inc., which is tightenedaround the circumference of the joined flanges 120-130. Deployment afterlaunch is accomplished by loosening the LightBand 140.

As illustrated in FIG. 1B, the individual spacecraft can be different insize and shape, provided they conform to an overall envelope and includethe appropriate flange 130. As illustrated in FIG. 1C, the adapter 100is designed to be situated concentric with the vertical axis 182 of thelaunch vehicle 180, and multiple adapters 100 can be stacked within thelaunch vehicle 180, again, for example, using a LightBand to couple theadapters 100.

Although the ESPA provides a standard architecture for designing thespacecraft interface to the launch vehicle and enables multiplespacecraft to be launched from a single launch vehicle, it is notparticularly efficient in volume and weight.

As illustrated in FIGS. 1B and 1C, a significant amount of volume iswasted between the vehicles 150, and between the vehicles 150 and thefairing 185 of the launch vehicle. In a typical configuration, thevolume efficiency (volume of spacecrafts/available payload volume) canbe as low as 50%.

Because the adapter 100 must support the attached spacecraft withminimal movement during launch, its weight can range from 400 pounds toover 600 pounds. The fairing 185 surrounding the spacecraft of eachadapter can amount to well over 1000 pounds. In a typical configuration,the mass efficiency (mass of spacecraft/launch vehicle capacity) rarelyexceeds 50% due to the mass of the adapter, fairing, and other‘overhead’ items.

Another problem with the ESPA architecture is the creation of convolutedload paths, as illustrated by the arrows 190 in FIG. 1C, which lead topotentially large deflections 195 that are difficult to prevent withoutadding substantial mass to the supporting elements. The potentiallylarge deflections 195 also necessitates constraints on the spacecraft'sdynamic envelope, to avoid collisions between spacecraft during launch.

Additionally, the ESPA architecture does not provide a means to controlthe deployment of the spacecraft. When the launch vehicle is in theproper deployment location, after the fairing has been released, thecoupling between the flanges 120-130 is released and the space vehicle‘tumbles away’ until its internal navigation and propulsion systemsdirect it to its proper station.

U.S. Pat. No. 5,522,569, issued 4 Jun. 1996, to Steffy et al. disclosesa “SATTELITE HAVING A STACKABLE CONFIGURATION” that provides mass andvolume efficiency, with a simplified load path. Relatively shortcylindrical satellites of the same diameter are stacked and bolted toeach other using three coupling devices arranged on the perimeter ofeach satellite. The bolts are secured using separation nuts that releasethe bolts for deployment; a spring mechanism in each coupling devicepropels the top-most satellite away from the stack. The bolting of eachsatellite to each other, and the lowermost satellite in the stack to thelaunch vehicles, provides a self-supporting structure with linearloading, with each set of couplers being preloaded (torqued) to minimizedeflection of the stack. However, this self-supporting structure isparticularly well suited for uniformly short cylindrical satellites, butif a tall satellite is included in the stack, the wall structure of thetall satellite would need to be sufficiently reinforced to avoidunwanted stack deflection.

U.S. Pat. No. 5,129,601, issued 14 Jul. 1992, to Henkel, discloses a“JACK SCREW PAYLOAD DEPLOYMENT SYSTEM” that uses a set of threemotor-driven screws on a baseplate that is attached to the launchvehicle. The motor-driven screws are threaded into attachment fittings(nuts) on the space vehicle to attach the space vehicle to thebaseplate. By un-screwing the screws at a predetermined speed, the spacevehicle can be ‘launched’ from the baseplate at a desired velocity.However, this arrangement is a single space vehicle deployment system,because once the space vehicle is screwed down to the baseplate, thescrews cannot be further rotated to accept other space vehicles. Ifmultiple space vehicles are threaded onto the screws sequentially, thescrews will again cease rotation when the lower space vehicle reachesthe baseplate, preventing the preloading (torqueing) of the upper spacevehicles, rendering the stack unstable for launch.

It would be advantageous to provide a system and method for installing,deploying, and recovering a plurality of spacecraft that provides anease of use and structural stability that is not currently available inexisting spacecraft deployment systems. It would be of further advantageto provide a system and method that supports a standardization ofspacecraft design that enables spacecraft from different sources to beefficiently arranged within the launch vehicle.

These advantages, and others, can be realized by defining a standard, orfamily of standard dimensions for the exterior shape of each spacecraft,with well defined placement of internal structures that facilitate thecoupling of multiple spacecraft in a stack above a baseplate that isconfigured to be fixedly attached to the launch vehicle.

In an embodiment of this invention, a plurality of threaded rods arearranged orthogonal to a surface of the baseplate, and each spacecraftincludes a channel through which each rod can traverse. Each spacecraftalso includes a coupling mechanism that selectively engages ordisengages each threaded rod. Each spacecraft is added to the stack byengaging its coupling mechanism while each of the preceding spacecraftdisengages its coupling mechanism. When the coupling mechanism isengaged and the threaded rods are rotated, the spacecraft travels alongthe threaded rods toward the baseplate; when the coupling mechanism isdisengaged, the threaded rod is free to rotate. When all of thespacecraft are added to the stack, the threaded rods engage a terminatorcomponent, which may be the nosecone of the launch vehicle, and arerotated while the coupling mechanisms in all of the spacecraft of thestack are disengaged. The threaded rods are screwed into the terminatorcomponent, thereby preloading the stack to the baseplate to withstandthe loads introduced during launch.

To deploy the spacecraft from the launch vehicle, the process isreversed. The terminator component is released by unscrewing thethreaded rods while the coupling mechanisms of all of the spacecraft aredisengaged. Upon release of the terminator component, the uppermostspacecraft in the stack engages its coupling mechanisms, therebyenabling the spacecraft to travel along the threaded rods, away from thebaseplate, when the threaded rods are further rotated. Each subsequentspacecraft is similarly ejected by engaging its coupling mechanisms asthe threaded rods are rotated.

The spacecraft deployment system of this invention may also be used toretrieve spacecraft, thereby reducing the amount of ‘space junk’ thatremains in orbit after the spacecraft has completed its mission. In suchan embodiment, the remainder portion of the launch vehicle with thebaseplate and threaded rods is directed to the spacecraft that is to beretrieved. The threaded rods enter the channels of the spacecraft, andthe spacecraft engages its coupling mechanisms to travel along thethreaded rods, toward the baseplate, when they are rotated. As eachsubsequent spacecraft is engaged, the spacecraft on the stack disengagetheir coupling mechanisms to enable the threaded rods to rotate. If thenosecone is available for retrieval, it is retrieved and used as theterminator component to preload the stack for re-entry; otherwise, thetopmost spacecraft may engage its coupling mechanism to serve as theterminator component.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention is explained in further detail, and by way of example,with reference to the accompanying drawings wherein:

FIGS. 1A-1C illustrate an example prior art spacecraft deploymentsystem.

FIG. 2 illustrates an example spacecraft deployment system using aspectsof this invention.

FIGS. 3A-3B illustrate an example rectilinear spacecraft using aspectsof this invention.

FIG. 4 illustrates an example cylindrical spacecraft using aspects ofthis invention.

FIGS. 5A-5B illustrate an example coupling mechanism using aspects ofthis invention.

FIG. 6 illustrates an example rack and pinion coupling mechanism usingaspects of this invention.

FIG. 7 illustrates another example embodiment of a coupling mechanismusing aspects of this invention.

FIGS. 8A-8B illustrate another example embodiment using aspects of thisinvention.

FIGS. 9A-9B illustrate another example embodiment using aspects of thisinvention.

FIG. 10 illustrates another example embodiment using aspects of thisinvention.

Throughout the drawings, the same reference numerals indicate similar orcorresponding features or functions. The drawings are included forillustrative purposes and are not intended to limit the scope of theinvention.

DETAILED DESCRIPTION

In the following description, for purposes of explanation rather thanlimitation, specific details are set forth such as the particulararchitecture, interfaces, techniques, etc., in order to provide athorough understanding of the concepts of the invention. However, itwill be apparent to those skilled in the art that the present inventionmay be practiced in other embodiments, which depart from these specificdetails. In like manner, the text of this description is directed to theexample embodiments as illustrated in the Figures, and is not intendedto limit the claimed invention beyond the limits expressly included inthe claims. For purposes of simplicity and clarity, detaileddescriptions of well-known devices, circuits, and methods are omitted soas not to obscure the description of the present invention withunnecessary detail.

Additionally, for purposes of explanation, the following terms are usedherein, with accompanying explanation. These explanations are providedfor ease of understanding, and are not intended to limit the claimedinvention beyond the limits expressly included in the claims.

Actuator: a mechanical device for moving or controlling another device.

Baseplate: a lowest structure in the stack of spacecraft, configured tobe fixedly attached to the launch vehicle.

Bolt: a rod with a helical thread.

Coupling mechanism: a structure that is able to selectively, join thespacecraft to a rod.

Launch vehicle: a rocket used to launch spacecraft.

Nosecone: a protective cone constituting the forward end of a launchvehicle.

Nut: a device with an internal thread that engages the thread of a bolt;as used herein, the nut need not completely encircle the bolt.

Pinion: a gear with teeth designed to mesh with a larger wheel or rack.

Preload: application of stress to a mechanical system; as used herein, acompression force to increase the rigidity of the stack of spacecraft.

Rack: a bar with teeth for operating with a pinion or worm gear totransform rotary motion to linear motion or vice versa.

Rod: a straight slender bar.

Stack of spacecraft: a plurality of spacecraft arranged vertically abovea baseplate.

Spacecraft: a vehicle or device designed for travel or operation outsidethe earth's atmosphere.

Terminator component: a topmost structure in the stack of spacecraft,configured to preload the stack to the baseplate.

Threaded rod: a rod with a plurality grooves or one or more helicalgrooves.

FIG. 2 illustrates an example embodiment of a spacecraft deploymentsystem that uses aspects of this invention. In this embodiment, a stack220 of spacecraft 221-229 is situated upon a baseplate 210 that isfixedly attached to the final stage 270 of a launch vehicle 280.Threaded rods 230 extend through channels (330 in FIGS. 3A-3B) in eachspacecraft 221-229, and into a terminator component 250, which isillustrated as the nosecone of the launch vehicle 280. One of skill inthe art will recognize that the terminator component 250 may be anycomponent that receives and captures the terminal end of the threadedrods 230. Optionally, protective elements, such as O-rings, EMI gaskets,etc. may be placed between the spacecraft 221-229.

Each spacecraft 221-229 includes coupling mechanisms 235 thatselectively engage or disengage the threaded rods 230. When the couplingmechanisms 235 of a spacecraft are engaged, the rotation of the threadedrods 230 cause the spacecraft to travel along the threaded rods 230,either toward the baseplate 210 for adding the spacecraft to the stack220, or away from the baseplate 210 for removing the spacecraft from thestack 220.

In the example of FIG. 2, each threaded rod 230 is driven by an actuator240 that selectively rotates the rod 230. Each actuator 240 of the stackis synchronized with each other actuator 240 so that the spacecraftmoves uniformly along each threaded rod 230 when its coupling mechanisms235 are engaged. The coupling mechanisms 235 of each spacecraft aresimilarly synchronized to uniformly engage or disengage the threadedrods 230.

As noted above, each spacecraft 221-229 is added to the stack bysituating the rods 230 into channels of the spacecraft and enabling thecoupling mechanisms of the spacecraft to engage the threaded rods 230.The channels may be flared at the bottom of the spacecraft to facilitatethe entry of the rods 230 into the channels. While the couplingmechanism 235 of the spacecraft to be added is engaged, the couplingmechanisms of all spacecraft currently on the stack are disengaged fromthe threaded rods 230, allowing the threaded rods 230 to rotate, therebypropelling the spacecraft along the threaded rods 230 toward thebaseplate 210.

Optionally, if the spacecraft is being added to the stack using a craneor other supporting mechanism, the spacecraft's coupling mechanisms maybe disengaged and the crane may lower the satellite onto the existingstack. Threading the spacecraft onto the existing stack is generallypreferred in order to avoid the complexities of a controlled uniformlowering of the spacecraft, particularly if the tolerance between therods and the channels of the spacecraft is tight. Threading thespacecraft onto the existing stack also allows the crane to remove themechanism to lift the spacecraft, before the spacecraft is in contactwith the stack.

By arranging the spacecraft 221-229 in a vertical stack, the load pathis along the rods 230 and the accompanying channels 330 in eachspacecraft. This vertical loading substantially reduces the complexityof load management as discussed above with respect to prior art ESPAsystem, as well as substantially reducing the potential deflection 195of the ESPA deployment system. Also as compared to the ESPA deploymentsystem, the volume consumed by each spacecraft is substantially reduced,as well as the volume required to accommodate the supporting adapter100. This reduction in volume allows for smaller and/or shorter launchvehicle fairings and/or larger spacecraft volume.

In preparation for launch, the stack is preloaded by rotating thethreaded rods into the terminator component at the top of the stack. Inan embodiment of this invention, the threaded rods may be ⅜″ steel rods,and are torqued to produce a preload of at least 5,000 lbf. to minimizemovement of the stack during launch. If the terminator component is inthe nosecone 250, the nosecone 250 may be configured to have a ‘lip’within which the fairing fits, with a corresponding lip at the top ofthe final stage 270. In this manner, panels of the fairing may besituated within the lip of the final stage 270, and extends into the lipof the nosecone 250 as the nosecone 250 is drawn down upon the threadedrods 230. When the nosecone 250 is subsequently unscrewed from thethreaded rods 230 during deployment, the panels of the fairing maysimply ‘fall away’ from the final stage 270.

To deploy the spacecraft 221-229 after launch, all of the couplingmechanisms of the spacecraft 221-229 are disengaged and the actuators240 are enabled to unscrew the threaded rods 230 from the terminatorcomponent 250. Thereafter, the uppermost spacecraft on the stack 220engages its coupling mechanism and the actuators are enabled to unscrewthe spacecraft from the stack. The deployment velocity of the spacecraftmay be precisely controlled by controlling the rotation speed of thethreaded rods 230, and higher velocities may be achieved compared toconventional spring-loaded deployments. Additionally, by projecting thespacecraft from the threaded rods 230, the likelihood or degree oftumbling is substantially reduced.

Although the spacecraft will typically be deployed individually,multiple spacecraft may be released in quick sequence by initiallycreating a space between the spacecraft, then simultaneously engagingeach of the coupling mechanisms of each of the multiple spacecraft.Alternatively, the multiple spacecraft could be deployed by sequencingthe engagement of each spacecraft without waiting for the upperspacecraft to be completely deployed from the threaded rods 230.

After deploying all of the spacecraft from the stack, the deploymentsystem may be used to retrieve other spacecraft by guiding the threadedrods 230 into the channels of the retrieved spacecraft and using thetechniques detailed above to create a stack of satellites atop thebaseplate 210. If the nosecone 250 is accessible, it can be used as theterminator component to secure the stack for re-entry; otherwise, theuppermost spacecraft may engage its coupling mechanism 235 to serve asthe terminator component.

FIGS. 3A-3B illustrate an example embodiment of a spacecraft 300 usingaspects of this invention. In this example, the spacecraft 300 issquare, with a channel 330 in each of the four corners of thespacecraft. Each channel 330 allows the threaded rod 230 to traversethrough the spacecraft 300. The channels 330 are structured to providesupport to the stack when the threaded rods 230 are attached to theterminator component and preloaded. The side panels 340 provide reactiveforces to shear forces during launch to minimize potential rotationaldeflection of the stack.

FIG. 3A illustrates coupling mechanisms 235′, 235″ that selectivelyengage or disengage the threaded rods 230. Although all of the couplingmechanisms 235 of a spacecraft would be concurrently engaged ordisengaged, the coupling mechanisms 235′ are illustrated in the engagedstate, while the coupling mechanisms 235″ are illustrated in thedisengaged state for ease of understanding. Although the channels 330are illustrated as being ‘closed’ structures, one of skill in the artwill recognize that the channels 330 may only partially enclose the rods230, to facilitate engagement by the coupling mechanisms 235.

Although the spacecraft of FIGS. 3A, 3B has a square shape with fourthreaded rods 230, one of skill in the art will recognize that othershapes may be used, with a corresponding increase or decrease in thenumber of threaded rods 230. If the spacecraft is a polyhedron, the rodswill generally be placed at the vertices. In general, the number of rodswill increase with the size and mass of the spacecraft, and the numberof sides will determine the volumetric efficiency by reducing the ‘emptyspace’ between the perimeter of the spacecraft and the fairing of thelaunch vehicle 280. However, a square shape may be preferred becausemost subsystem components of a spacecraft are rectangular.

One of skill in the art will also recognize that additional rods thatare only threaded to couple to the terminator component 250 may be usedfor preloading the stack. In such a configuration, the threaded rods 230may be used only to transfer the spacecraft onto or off the stack 220.

If only one threaded rod 230 is used, it may be preferably located atthe center of the stack of spacecraft 220. This single threaded rodconfiguration has the advantage that each spacecraft need only have onecoupling mechanism 235, and the need to synchronize multiple couplingmechanisms 235 in the spacecraft is eliminated. If the threaded rod 230is at the center of the stack, however, the coupling mechanism 235 andspace for the threaded rod 230 need to be in the center of thespacecraft, which may complicate the configuration of themission-specific components in the spacecraft.

FIG. 4 illustrates an alternative embodiment, wherein the spacecraft 400is cylindrical, and may form the fairing of the launch vehicle 280,thereby optimizing the available volume. In this example, thecylindrical walls of the spacecraft provide the channels (notillustrated) for the threaded rods 230, or for additional rods that areonly threaded to couple to the terminator component 250. The threadedrods 230 are preferably arranged symmetrically about the perimeter ofthe spacecraft, as are the additional rods, if any.

FIGS. 5A-5B each illustrate top and profile views of an example couplingmechanism 235 comprising a device 510 that selectively extends (FIG. 5B)and withdraws (FIG. 5A) a receptor 520 to selectively engage anddisengage the threaded rod 230. In this example embodiment, the receptor520 includes one or more tongs 525 that are structured to engage thethreads 232 of the threaded rod 230. Other means of selectively engagingthe threads 232 of the threaded rod 230 will be evident to one of skillin the art, include split nuts and the like. Although not illustrated inFIGS. 5A-5B, the threads 232 of the threaded rod 230 are preferablytrapezoidal in shape to facilitate the engagement of the tongs 525 orthe internal threads of a split nut.

FIG. 6 illustrates another example embodiment of a coupling mechanism235 that uses a rack and pinion arrangement to selectively engage anddisengage the threaded rod 230. The axle 615 of the pinion 610 isattached to a coupling mechanism 620 that is attached to the spacecraft.In a simple embodiment, the coupling mechanism 620 selectively inhibitsor releases the rotation of the pinion 610, such as a controlled brake.When the rotation of the pinion 610 is released, the threaded rod 230 isfree to turn without exerting any lateral forces on the axle 615 of thepinion 610. When the rotation of the pinion 610 is inhibited, therotation of the threaded rod 230 exerts a vertical force on the pinion610, causing the mechanism 620, and the spacecraft to move vertically,either up or down depending on the direction of the rotation of thethreaded rod 230.

FIG. 7 illustrates an alternative embodiment wherein the threaded rod730 includes a conventional helical thread 732 to screw into theterminator component 150, and a series of individual horizontal threads734, or slots below the terminator component 150. In such an embodiment,the coupling mechanism 720 includes a motor (not illustrated) thatserves as an actuator that rotates the pinion 610, causing thespacecraft to travel along the slots of the threaded rod 730; and therotation of the threaded rod 730 by the actuator (240 of FIG. 2) is onlyused to attach or release the terminator component 150. That is, in thisembodiment two independent actuators are used: an actuator 240 thatcontrols the preloading (the ‘securing’ actuator), and an actuator 720(the ‘locating’ actuator) that controls the travel of the spacecraftalong the threaded rod 730.

In this embodiment, the actuator 720 controls the engagement anddisengagement of the coupling mechanism 610, as well as the travel alongthe treaded rod 730. As the spacecraft is added to the stack, theactuator 720 rotates the pinion 610, which exerts a vertical force on(engages) the threaded rod 730. When the spacecraft is on the stack, theactuator 720 removes the rotation force on the pinion 610, therebyremoving the vertical force on (disengages) the threaded rod 730,allowing the threaded rod 730 to be rotated by the actuator 240.

Optionally, the actuators 720 of the spacecraft on the stack may beengaged to reapply the vertical force on the threaded rod after thethreaded rod 230 applies the preload to the stack, to further enhancethe rigidity of the stack.

One of skill in the art will recognize that the rod 730 may comprise ahollow outer cylinder that includes the threads 834 and slots forengagement by the actuator 720, and a concentric separate rod thatextends from the actuator 240 to the threaded end 732 that screws intothe terminal component 150. In this manner, the rotation of the threadedrod 730 will be independent of the vertical forces placed on the outercylinder by the pinion 610.

One of skill in the art will also recognize that different features ofthis invention may be combined. For example, the coupling mechanism 620of FIG. 6 may also include a motor that rotates the pinion 610, to beused in conjunction with, or independent of, the rotation of thethreaded rod 230.

FIGS. 8A-8B illustrate another alternative embodiment wherein themovement along the rods is provided by motorized mechanisms in eachspacecraft. As in FIG. 7, in FIG. 8A, a pinion 610 is driven by anactuator 720, but in this embodiment, the rod is a ‘rack’ 830 upon whichthe pinion 610 travels.

As illustrated in FIG. 8B, additional rods 832 may be used to align thespacecraft 800 as it travels along the racks 830. In this embodiment,the rods 832 may be smooth in the region of spacecraft locations, withthreads that engage the terminal component 150 (not shown).

In some embodiments, as illustrated in FIG. 8B, the rods 830 may beexternal to the spacecraft, rather than in channels (330 of FIG. 3)within the spacecraft. In some embodiments, the channels of thespacecraft could be “C” channels 835 within which the rods 830 aresituated. In some embodiments, the rods 830 may be affixed to thefairing 185 of the launch vehicle.

FIGS. 9A-9B illustrated another example embodiment wherein eachspacecraft 900 includes a plurality of motors 910 that propel eachspacecraft along the threaded rods 230. The axis of the motors 910 areconcentric about their corresponding threaded rods 230, and the rotor ofeach motor 910 includes threads that engage the threaded rods 230. Inthis manner, a rotation of the rotor causes the motor 910, and theattached spacecraft 900, to travel along the threaded rod 230.

If the threaded rods 230 are used to secure the terminal component 150via a rotation of each threaded rod, the motors 910 will include acoupling mechanism that selectively engages and disengages the threadedrod to enable rotation of the threaded rod. Alternatively, the terminalcomponent may include a motor 910 that secures the terminal component tothe stack.

One of skill in the art will recognize that the spacecraft will includea means of communicating with the launch vehicle (or launch control),and in some instances, may include a means of communicating with theother spacecraft in the stack. In the pre-flight stage, thiscommunication may be via external ports on each spacecraft, but duringand after launch, the stack of satellites must operate as a cohesiveunit, if for no other reason than synchronizing the discharge of thesatellites from the stack.

In embodiments of this invention, each spacecraft includes one or moreports 850 at both the ‘top’ and ‘bottom’ surfaces that contact thespacecraft above and below it on the stack, respectively. The ports onthe lowermost spacecraft on the stack are coupled to ports on thebaseplate, which preferably includes a controller that monitors andreports the status of each satellite to launch control, and controls thedischarge of each satellite from the stack. Alternatively, the lowermostspacecraft may be configured to perform this control function.

The ports are preferably situated near the channels that contain thethreaded rods 230, to assure connectivity between the ports. These portsshould be self-aligning, such that locating the spacecraft on the stack,or on the baseplate, automatically provides connectivity. These portsmay provide the communication of other-than-electrical elements, such asliquids or gasses, between and among the spacecraft.

Each spacecraft may also be configured to determine, and optionallyreport, its location along the rods. Such determination may be made, forexample, to distinguish between the spacecraft actually contacting thelower spacecraft or baseplate and the spacecraft encountering unexpectedresistance (binding) as it travels along the rod. The location along therod may also be used during the deployment of each satellite, todetermine and/or control the velocity at which the spacecraft isdeployed. In some embodiments, the location of the spacecraft may beused to synchronize or activate an operation of the launch vehicle, anoperation of the controller that deploys the spacecraft, or an operationof the spacecraft itself, such as a determination of ignition of jets onthe spacecraft.

In a simple embodiment, if the spacecraft is moved along the rod via apinion or other gear arrangement, the number of rotations of the gearfrom an initial starting location may be used to determine how far thespacecraft has traveled along the rod.

In a more complex embodiment, the rod may include optical or mechanicalmarkings that delineate locations along the rod, and the spacecraft mayinclude an optical or mechanical detector, or both, that reads anddecodes the markings. For example, the threads on the rod may include‘flats’, and the spacecraft may include a cam arrangement that engages acounter as each flat is encountered. Electronic location detectiontechniques may also be used, wherein at different locations along therod, a different electrical signal is received by a detector in thespacecraft; or, the spacecraft emits a signal and a controller in thebaseplate determines the location of the spacecraft based on propagationcharacteristics of the rod.

A combination of techniques may also be used. For example, the rod mayinclude visual markings at fixed intervals, with mechanical featuresthat enable determining the location relative to the visual markers fora finer location determination.

The foregoing merely illustrates the principles of the invention. Itwill thus be appreciated that those skilled in the art will be able todevise various arrangements which, although not explicitly described orshown herein, embody the principles of the invention and are thus withinits spirit and scope.

For example, although the structure of the rods has been disclosed as asolid rod with threads, one of skill in the art will recognize thatalternative structures may be used. A hollow rod, for example, mayprovide a higher specific stiffness, or a given stiffness with lessmass, than a solid rod. A hollow rod may also provide lower stress andwear on the actuators and/or higher rotational rates.

Although a rod with grooves is preferred for locating the spacecraft onthe stack, one of skill in the art will recognize that a smooth rod witha helical screw on the end for coupling with the terminal element may beused if the spacecraft is equipped with a coupling mechanism that canprovide sufficient surface friction to enable the spacecraft to travelalong the smooth rod when the coupling mechanism is engaged, such as awheel with a rubber perimeter in place of the aforementioned pinion.

One of skill in the art will also recognize that the baseplate that isattached to the launch vehicle may be incorporated within the lowermostsatellite on the stack. In like manner, the terminator component may beincorporated within the uppermost satellite on the stack. In such anembodiment, the coupling mechanism of the uppermost satellite may serveas the terminator component.

Although this invention is presented as a means for achieving volumeefficiency by standardizing the size of the satellites 220 to fill theavailable volume more efficiently than conventional multiple-payloadspacecraft, the principles of this invention may be applied for otherstandard configurations to address unique requirements. For example,most satellites include solar panels that are ‘unfolded’ when thesatellite is deployed, and this unfolding may include multiple hingedelements. This unfolding presents design challenges, including, forexample, assuring that the satellite has an internal energy supply topower the unfolding apparatus until the panels are deployed to generateelectricity. Additionally, the satellite could be totally disabled ifthere is a mechanical or electrical malfunction that affects theunfolding process.

FIG. 10 illustrates an example embodiment 100 wherein the solar panels1030 are not folded, but are simply hinged 1025 to the satellites 1020.Multiple (typically 4) panels 1030 may be situated about the perimeterof the satellites 1020. Conventionally, the allocated volume allocatedto a spacecraft is bounded by the planes 1040 that are defined by theupper and lower surfaces of the satellite 1020. In the embodiment ofFIG. 10, however, the solar panels 1030 extend beyond the plane 1040 ofthe bottom surface of the satellite 1020, into the volume that isconventionally allocated to the lower satellite, or group of satellites,or the volume adjacent the fuel tank 1010.

Even though the solar panels 1030 will typically need to be deployedalong a common plane that faces the sun, this deployment only requiresthe rotation of the solar panels 1030 about the hinge 1025 that couplesit to the spacecraft 1020. If the hinge 1025 is spring-loaded, andlatched in a tensioned state, this deployment can be accomplished by amere release of the latch when the satellite 1020 is released from thespacecraft 1000.

Although the satellites 1020 are illustrated as being similar to eachother, as would be common in a ‘constellation’ of satellites, one ofskill in the art will recognize that different sized satellites may beincluded in the same spacecraft 1000. For example, the uppermostsatellites may be satellites that do not have these hinged solar panel,and accordingly may be sized to occupy more of the volume of thespacecraft 1000 within the conventional space defined by the planes 1040between each spacecraft. In like manner, if the spacecraft 1000 does notallow the solar panels 1030 to extend into the volume allocated to thefuel tank 1010, the lowermost satellites may be satellites without solarpanels 1030.

These and other system configuration and optimization features will beevident to one of ordinary skill in the art in view of this disclosure,and are included within the scope of the following claims.

In interpreting these claims, it should be understood that:

a) the word “comprising” does not exclude the presence of other elementsor acts than those listed in a given claim;

b) the word “a” or “an” preceding an element does not exclude thepresence of a plurality of such elements;

c) any reference signs in the claims do not limit their scope;

d) several “means” may be represented by the same item or hardware orsoftware implemented structure or function;

e) any of the disclosed devices or portions thereof may be combinedtogether or separated into further portions unless specifically statedotherwise;

f) no specific sequence of acts is intended to be required unlessspecifically indicated; and

g) the term “plurality of” an element includes two or more of theclaimed element, and does not imply any particular range of number ofelements; that is, a plurality of elements can be as few as twoelements, and can include an immeasurable number of elements.

We claim:
 1. A spacecraft comprising: a body; wherein the body comprisesat least one channel through which at least one deployment element of adeployment system traverses the spacecraft; wherein the deploymentsystem is configured to deploy the spacecraft into space after launchvia the deployment element; at least one coupling mechanism; wherein thecoupling mechanism is configured to selectively engage the deploymentelement; wherein an engagement of the deployment element by the couplingmechanism causes the spacecraft to travel in a direction parallel to alongitudinal direction of the deployment element, wherein adisengagement of the deployment element by the coupling mechanism causesthe spacecraft to be substantially independent of the operation of thedeployment system.
 2. The spacecraft of claim 1, wherein the deploymentelement is a threaded rod, and wherein the coupling mechanismselectively extends a receptor to engage the threaded rod, and withdrawsthe receptor to disengage the threaded rod.
 3. The spacecraft of claim1, wherein the deployment element is a threaded rod, and wherein thecoupling mechanism comprises a rack and pinion arrangement; wherein thepinion comprises teeth that extend into threads of the threaded rod,wherein the coupling mechanism selectively inhibits rotation of thepinion to engage the threaded rod, and releases the pinion to freelyrotate to disengage the threaded rod.
 4. The spacecraft of claim 1,wherein the deployment element is a rack with horizontal slots, andwherein the coupling mechanism comprises a motor that rotates a pinion;wherein the pinion comprises teeth that extend into threads of thethreaded rod, wherein the coupling mechanism selectively activates themotor to engage the rack and pinion by rotating the pinion to exert aforce on the rack, causing the spacecraft to travel in the directionparallel to the longitudinal direction of the rack, and wherein thecoupling mechanism disengages the rack and pinion by ceasing therotation of the pinion, thereby releasing the force on the rack.
 5. Thespacecraft of claim 1, wherein the spacecraft is configured to join astack of other spacecraft by an alignment of the channel of thespacecraft with other channels of the other spacecraft, thereby enablingthe deployment element to traverse the stack of spacecraft.
 6. Thespacecraft of claim 5, wherein the coupling mechanism selectivelyengages and disengages the deployment element to facilitate situatingthe spacecraft on the stack.
 7. The spacecraft of claim 5, wherein thespacecraft includes at least one communication port that is configuredto couple to ports of adjacent spacecraft on the stack.
 8. Thespacecraft of claim 7, wherein the coupled communication ports on thestack enable communication between the spacecraft and the deploymentsystem.
 9. The spacecraft of claim 8, wherein the communication ports onthe stack are also configured to communicate power from the deploymentsystem to the spacecraft.
 10. The spacecraft of claim 1, wherein thespacecraft comprises a communication system that enables communicationbetween the spacecraft and the deployment system.
 11. The spacecraft ofclaim 10, wherein the spacecraft comprises one or more sensors thatsense a state of the spacecraft relative to the deployment element, andwherein the communication system communicates the state of thespacecraft to the deployment system.
 12. The spacecraft of claim 1,wherein the body is rectilinear.
 13. The spacecraft of claim 1, whereinthe body is cylindrical.
 14. The spacecraft of claim 13, wherein anouter wall of the body forms at least a portion of a fairing of thelaunch vehicle.
 15. A spacecraft deployment system comprising: adeployment controller; at least one deployment element; and at least oneactivator; wherein the deployment controller is configured to controlthe actuator to deploy a plurality of spacecraft into space after launchof the plurality of spacecraft; wherein the deployment element isconfigured to be selectively engaged by a coupling mechanism of eachspacecraft of the plurality of spacecraft; wherein the deploymentcontroller causes the activator to cause each spacecraft to travel in adirection parallel to a longitudinal direction of the deployment elementwhen the deployment element is engaged by the coupling mechanism of eachspacecraft.
 16. The spacecraft deployment system of claim 15, whereinthe deployment controller is configured to control the couplingmechanism of each spacecraft to selectively engage and disengage thecoupling mechanism from the coupling element.
 17. The spacecraftdeployment system of claim 15, wherein the deployment controller isconfigured to control the actuator to facilitate placement of eachspacecraft on a stack of spacecraft in preparation for launching theplurality of spacecraft.
 18. The spacecraft deployment system of claim17, wherein the deployment element is a threaded rod.
 19. The spacecraftdeployment system of claim 18, wherein the deployment controller isconfigured to control the actuator to engage the deployment element witha terminal component and to torque the threaded rod to provide a preloadto the stack.
 20. The spacecraft deployment system of claim 19, whereinthe deployment controller is configured to disengage the couplingmechanism of each spacecraft in the stack while the actuator torques thethreaded rod into the terminal component.
 21. The spacecraft deploymentsystem of claim 15, wherein the spacecraft deployment system comprises anavigation system and a maneuvering system to facilitate deployment ofeach spacecraft at its intended location in space.
 22. The spacecraftdeployment system of claim 21, wherein the spacecraft deployment systemuses the navigation and maneuvering systems to retrieve other spacecraftfrom space via the deployment controller, actuator, and deploymentelement.
 23. The spacecraft deployment system of claim 22, wherein thedeployment element is a threaded rod.
 24. The spacecraft of claim 15,wherein the deployment element is a hollow rod.